Operating efficiencies of air-breathing turbine engines have historically been altered by variations in the position of the exit edges of stator vanes, thereby controlling and adjusting the flow of exit gases impinging upon rotor vanes (customarily referred to as "buckets"). Most modern airplane turbine engines typically incorporate from one to six rings or stages of variable stator vanes in the compression stage, each of which may be comprised of approximately sixty to seventy variable vanes per ring. It has been known for decades that controlling the deflection of the individual stator vanes comprising each ring to correlate gas flow through the engine to the flight regime of the aircraft is beneficial, particularly in respect of fuel efficiency. More specifically, varying the space or equivalent aperture dimension between adjacent stators with variations in the fuel demanded for work performance requirements during or preparatory to flight is highly advantageous. Among those patent references which may be conferred with these thoughts in mind may be mentioned U.S. Pat. Nos. 2,065,974, 2,219,994 and 2,676,458.
With nearly fifty years' recognition of the advantages to be obtained by controlled manipulation or orientation of stator vanes, state of the art turbine engine are a vastly complicated affair of hardware as respects vane actuation systems. Looking to aircraft engines of the most recent technical evolution, variable stator vane actuation systems commonly include a plurality of hydraulic actuators, actuation mechanisms and linkages, and flexible feedback cables to achieve these results. Each variable stator vane actuator is usually connected through a clevis link and a multiple stage bellcrank to a master rod. Adjustable linkages interconnect variable vane actuation rings to those bell cranks which are, in turn, associated with the master rod. Connections between the actuator, clevis links, and master rod are made with bolts and bushings for stability; while other linkages are made through bolts and uniballs to eliminate misalignment or binding. The actuation rings, which are conventionally connected at a horizontal split-line of the compressor casing, rotate circumferentially about a horizontal axis of the compressor; whereby movement of the rings is transmitted to the individual vanes through vane actuating levers. A flexible cable attached to the linkage transmits a feedback signal to the main engine control. The feedback mechanism in the control repositions a pilot valve to stabilize the actuator signal when the vanes achieve the scheduled position. The variable stator vanes are positioned by variable stator vane actuators which are operated by fuel pressure from the main engine control. Within the main engine control is a variable stator vane scheduling cam having a three dimensional profile, which is positioned by engine speed and compressor inlet temperature signals. A variable stator vane feedback mechanism transmits actual vane position to the control while a variable stator vane pilot valve is positioned as a result of the comparison of the scheduling cam position and the feedback signal. Changes in engine speed rotate the scheduling cam while changes in compressor inlet temperature translate the cam axially. In turn, movement of the cam reposition the pilot valve which ports high pressure to either the head-end (closing) or rod-end (opening) of the variable stator vane actuators while venting the other end to bypass pressure.
As can be appreciated from the very terse description aforesaid, modern aircraft turbine engines incorporate very complicated mechanical mechanisms to achieve this required end of variation in stator positioning. There are substantial tradeoffs against the advantage of improved fuel efficiency when employing these types of assemblies. For example, initial adjustment, which must be made within carefully controlled limits, is a tedious and highly labor-itensive task. There is also a fair weight penalty associated with these elaborate mechanisms.
Within the context of the present invention, there have been various prior proposals with an eye toward the development of a so-called "smart" variable stator vane. These approaches have sought to take advantage of bimetallic construction and analogous geometric profiles so that the variations in inlet temperature of the working fluid may be used beneficially to alter the profile or camber of the stator. Exemplary of patented approaches for the automatic control over variations in the shape of stator vane in a gas turbine engine responsive to temperature changes in the environment surrounding those vanes are U.S. Pat. Nos. 3,038,698 and 3,042,371. Each of the proposals disclosed in those patents relies upon a conventional bimetallic couple. As the inlet temperature changes, so too does the chamber of the vanes. However, these approachs suffer inherent and serious limiting disadvantages all but excluding the use of such constructions in modern aircraft turbine engines. For example, the vanes disclosed in those references show mechanically unsupported elements which are likely to experience early failure in the noisy vibrational environment of a turbine engine. Both also overlook the distinct susceptibility of bimetallic couples to fatigue failure because of the high stress riser at the junction of the strip elements--the face-to-face juncture being prone to separation. Yet a worse reliability problem faces these types of proposals due to joining along vane edges by welding. As a rule of thumb, welded structures are typically avoided in environments which are subject to vibration as the predicted mode of failure is via cracking.
There are, of course, other thermally-responsive metallic compositions known in the art. "Nitinol", remarkable alloys of titanium, nickel and as of recently such other alloying constituents as cobalt and iron, to name but a few, exhibit the curious metallurgical phenomenon of thermal memory. Nitinol which has been formed at relatively high temperature may subsequently be crumpled, bent, or otherwise deformed at room temperature; then, due to its memory for the preformed shape, return thereto upon the application of a small amount of heat. Somewhat representative of these types of compositions are U.S. Pat. Nos. 3,764,227, 3,403,238 and 3,558,369. Thermal memory to the contrary notwithstanding, these types of compositions remain laboratory curiosities more than engineering alloys. There is currently no practical means of adapting this variety of composition to a controllable, repetitive deflection as is demanded by variations in stator vane camber within an aircraft engine.
Other references of incidental interest include U.S. Pat. Nos. 2,114,567 and 3,930,626, British Pat. No. 947,118 and German Pat. No. 1528887. While generally pertinent to temperature-responsive machinery elements, none adds much to those discussed in greater detail above.
Given the current state of the art, the need exists to provide a vane assembly, and especially a stator vane assembly for an aircraft turbine engine, which overcomes the tremendous complexity of current design on the one hand and eliminates the lack of reliability of simpler approaches on the other hand. Replacing complicated, costly, heavy actuator mechanisms without sacrificing reliability or efficiency in use is a goal heretofore elusive following the direction of the prior art.